This invention relates to the structure of a liquid-fueled rocket engine, and, more particularly, to the structure at the injector end of the combustion chamber.
A typical liquid-fueled rocket engine includes a generally cylindrical combustion chamber, with an injector attached to its injector end and a flared nozzle attached to its nozzle end. A liquid propellant including fuel and an oxidizer flows through injector ports in the injector and into the combustion chamber. The propellant mixes and ignites in the combustion chamber. The hot gas resulting from the combustion expands through the nozzle and drives the rocket engine and the attached rocket structure in the direction opposite to that in which the nozzle is pointed.
In order to achieve maximum efficiency of the rocket engine in terms of thrust per weight of propellant and maximum specific impulse, the mixing of the propellant components must be complete to encourage complete combustion. An internal step structure at the injector end of the combustion chamber had been previously developed to promote the complete mixing and combustion. However, early versions of the rocket engine having the internal step structure within the combustion chamber suffered from two problems. First, their efficiency, while better than that of a rocket engine of comparable weight but having no internal step structure, was not as good as desired. Second, in some cases there was premature failure of the wall of the combustion chamber during fabrication or during test firings.
The second problem is of particular concern, because a preferred embodiment of the rocket engine is to be used in commercial communications satellites. The rocket engine boosts the satellite from low-earth orbit to geosynchronous orbit or initiates interplanetary missions. If the rocket engine combustion chamber fails prior to completion of its mission, the entire satellite may be lost to service because it does not achieve the required geosynchronous orbit.
Thus, while the rocket engine with internal step structure in the combustion chamber offers important potential benefits in terms of improved efficiency, those benefits have not been fully realized in initial forms of the rocket engine. There is a need for a design which achieves maximum efficiency with high reliability and low likelihood of premature failure. The present invention fulfills, this need, and further provides related advantages.